1. Field of the Invention
The present invention concerns a method of placing a geostationary telecommunication satellite in orbit.
2. Description of the Prior Art
A telecommunication satellite is conventionally placed in its orbit (during the "pre-operational" phase) by a method comprising the following stages:
The launch vehicle injects the satellite into a transfer orbit; this intermediate orbit is elliptical and the launch vehicle injects the satellite into it substantially at the perigee.
Acquisition of the attitude for the apogee burn or maneuver; the apogee burn is carried out near the apogee of the transfer orbit in order to place the satellite in its final orbit which is a circular orbit with a radius substantially equal to the radius at the apogee of the transfer orbit.
The apogee burn or maneuver is performed near the apogee of the transfer orbit by firing the apogee thruster of the satellite.
When the final circular orbit is achieved (after one or more apogee maneuvers) the satellite "drifts" in this orbit until it reaches its final geostationary orbital position. It is not until this stage that the payload is rendered operational and the "operational phase" of the life of the satellite begins.
During the operational phase most current telecommunication satellites are stabilized relative to three axes by a fixed wheel providing a gyroscope function which enables indirect control of yaw by way of roll attitude control. Pitch control is achieved by varying the speed of the wheel. This minimal system provides a pointing accuracy which depends on the sensor used, on the value of the kinetic moment and on the resolution of the control torque. The accuracy may be improved by using supplementary devices such as magnetic coils, solar panels, inclined wheels or transverse reaction wheels, but this is to the detriment of weight and cost.
Most geostationary telecommunication satellites currently in orbit have a so-called north-south configuration in which the entire satellite, in other words the platform and the payload, is pointed permanently towards the Earth, the solar power generator being deployed and oriented along the axis perpendicular to the orbit and the solar panels turning about their longitudinal axis so as to point towards the sun at all times. In a configuration of this kind the axis of the momentum wheel is perpendicular to the thrust vector when the apogee thruster is fired. The momentum wheel must therefore not be rotating during the apogee maneuver.
There is described in the literature, for example in the document FR-A-2472509, another configuration which uses a satellite platform and solar panels pointed towards the sun at all times and a payload which is rotatably mounted on the platform so that it points towards the Earth at all times. This requires that the payload be entirely disposed on a north or south face perpendicular to the launch vehicle axis in the launch configuration. In this case the axis of the momentum wheel is aligned with the thrust imparted by the apogee thruster.
With the first of these configurations, which is that of most geostationary satellites currently in orbit, there are two main methods for stabilizing the satellite when it is moved from the transfer orbit to the final orbit.
The so-called "spin transfer" method entails injecting the satellite into its transfer orbit in such a way that it rotates (spins) about the axis of the apogee thruster. The successive phases of this method are as follows, for example:
the satellite is injected by the launch vehicle into a rotating attitude with the following characteristics:
attitude precision less than 6.degree.,
transverse angular speed less than 2.degree./second,
spin speed approximately 5 revolutions/minute,
the stabilizing system is powered up, the antennas are deployed and the spin speed is increased to 13 revolutions/minute,
the attitude and the spin speed are determined by telemetry using the terrestrial and solar elevation sensor,
the spin axis is reoriented before each apogee maneuver,
one or more apogee maneuvers are performed,
the spin speed is reduced,
the solar acquisition is performed and the solar generators are deployed,
the final attitude is acquired.
A method of this kind, which is very secure given the rotation imparted to the satellite when it is moved from the transfer orbit to its final orbit, has the advantage of minimizing the attitude maneuvers and the equipment required for final orbit injection. On the other hand, the effects of liquid tossing (fuel and combustion-supporting liquid) are difficult to control and model given the large quantities of propellants on the satellite. Also, this "spin" configuration has the disadvantage of imposing constraints regarding the inertia ratio and problems with laying out and balancing the satellite to meet the stability criteria.
The so-called "three-axis stabilized transfer method" entails injecting the satellite into its transfer orbit with a three-axis stabilized attitude and using dedicated equipment to control orientation throughout the transfer. The successive phases of a method of this kind are, for example:
the satellite is injected by its launch vehicle into a three-axis attitude with an attitude precision in the order of 3.degree., virtually no angular speed in the required attitude and with the solar generator facing towards the sun,
calibration of rate gyro drift, solar acquisition, terrestrial acquisition, rotation through 360.degree. about each axis for fine calibration of the rate gyros,
the solar generator is pointed towards the sun,
the rate gyros are calibrated again (having drifted in the meantime),
the apogee burn attitude is acquired,
the apogee maneuver is performed (one or more times), controlled by the rate gyros,
sun acquisition and deployment of the solar generator,
nominal attitude acquisition.
This second method has the disadvantage of not offering the inherent stabilization security of the spin transfer method (in the event of an equipment failure, for example). Also, it requires equipment specific to transfer of the satellite from its transfer orbit to its final orbit, including:
a sophisticated solar sensor with multiple optical heads,
an infra red terrestrial sensor, and
a three-axis integrating rate gyro assembly.
As a result it is complex and costly.
Another disadvantage of these known methods is that they are not autonomous and require significant ground support.
The invention is directed to remedying these disadvantages.